In a turbomachine, such as a gas turbine engine, air is pressurized in a compressor then mixed with fuel and burned in a combustor to generate hot combustion gases. The hot combustion gases are expanded within a gas turbine of the engine where energy is extracted to power the compressor and to produce useful work, such turning a generator to produce electricity. The hot combustion gases travel through a series of turbine stages. A turbine stage may include a row of stationary vanes followed by a row of rotating turbine blades, where the turbine blades extract energy from the hot combustion gases for powering the compressor and providing output power. Since the turbine blades are directly exposed to the hot combustion gases, they are typically provided with internal cooling circuits which channel a coolant, such as compressor bleed air, through the airfoil of the blade and through various film cooling holes around the surface thereof. One type of airfoil extends from a root at a blade platform, which defines the radially inner flow path for the combustion gas, to a radially outer cap or blade tip section, and includes opposite pressure and suction sides extending axially from leading to trailing edges of the airfoil. The cooling circuit extends inside the airfoil between the pressure and suction sides and is bounded at its top by the blade tip section.
The gas turbine engine efficiency is, at least in part, dependent upon the extent to which the high temperature gases leak across the gap between the turbine blade tips and the seals or shrouds which surround them. The leakage quantity is typically minimized by positioning the radially-outward blade tip section in close proximity to the outer air seal. However, differential thermal elongation and dynamic forces between the blade tip section and outer air seal can cause rubbing therebetween. Also, it should be noted that the heat load on the turbine blade tip section is a function of leakage flow over the blade tip section. Specifically, a high leakage flow will induce a high heat load to the blade tip section, such that gas leakage across the blade tip section and cooling of the blade tip section have to be addressed as a single problem. In a typical construction, see FIG. 9, the blade tip section 204 of an airfoil 200 has been provided with a squealer tip rail 202 extending radially outwardly a short distance from the blade tip section 204, and extending substantially completely around the perimeter of the airfoil 200 to define an inner squealer tip pocket 206 facing radially outwardly. The squealer tip rail 202 is provided for spacing radially closely adjacent to the stationary outer seal wall, or outer turbine shroud, to provide a relatively small clearance gap therebetween to seal or restrict the flow of gas across the blade tip section 204.
The squealer tip rail 202 is a solid metal projection of the airfoil 200, and is directly heated by the combustion gas which flows thereover, as illustrated by flow lines 208. In addition, a vortex flow 210 of hot gases may be formed on the suction side of the airfoil 200 adjacent the blade tip. The squealer tip rail 202 is cooled by a cooling fluid, such as air, channeled from an airfoil cooling circuit to the blade tip section 204 to convect heat away from the area of the squealer tip pocket 206. Convective cooling holes 214 may be provided in the squealer tip pocket 206 located along the squealer tip rail 202, as illustrated in FIG. 9. In addition, heat from the squealer tip rail 202 may be conducted into the squealer tip section 204 and convected away internally of the airfoil 200 by the cooling fluid channeled through the internal cooling circuit. The squealer tip section 204, including the squealer tip rail 202, typically operates at temperatures above that of the remainder of the airfoil 200 and can be a life limiting element of the airfoil 200 in a hot turbine environment. In particular, it is known in the art that the portion of the airfoil 200 located at the intersection of the pressure side airfoil surface 218 and the blade tip section 204 is subject to very high heat loads and accordingly is more likely to experience thermal distress.
Cooling to the pressure side airfoil surface 218 may be provided by a row of film cooling holes 216 located on the pressure side of the airfoil outer wall, extending from the leading edge to the trailing edge of the airfoil 200, immediately below the blade tip section 204 for providing a cooling fluid film which flows upwardly over the pressure side of the airfoil 200.